This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine airfoil having an auxiliary flow channel for receiving and communicating a portion of core airflow through the airfoil.
Gas turbine engines typically include at least a compressor section, a combustor section, and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section as well as other gas turbine engine loads.
The compressor and turbine sections typically include alternating rows of rotating blades and flow directing vanes. In the turbine section, the rotating blades extract energy from the core airflow that is communicated through the gas turbine engine, while the vanes direct the core airflow to a downstream row of blades.
The vanes can be manufactured to a fixed flow area that is optimized for a single flight point. Alternatively, it is possible to alter the flow area (i.e., cascade channel) between two adjacent vanes by providing one or more variable vanes that rotate about a given axis. Altering the flow area in this manner can expose downstream components to incidence angle variation. For example, rotating the variable vanes may alter the incidence angle at which hot combustion gases impinge upon rotor blades located downstream from the variable vanes, thereby potentially moving the flow stagnation point to a non-optimal location.